1. Field of the Invention
The present invention relates to forming advanced composite materials. More particularly, the present invention relates to a process for creating a single piece co-cured composite structure.
2. Technical Background
Advanced composite materials have presented a promising alternative to metals, plastics, and fiberglass. Advanced composite materials have the advantage of being very lightweight and have a high strength. Particularly, carbon-epoxy composite material are among the most promising of the advanced composite materials. Composite materials may also be formed into any number of different shapes, depending upon the application. Because of these advantages, composite materials are being employed in many different fields, such as aeronautics.
However, composite materials also have some disadvantages. Composite materials can be several times more expensive than metals or fiberglass. Despite the desirable strength and weight characteristics of composite materials, the price of the composite materials can often be cost prohibitive.
Additionally, forming composite materials into a desired structure can require complex assembly equipment and multiple manufacturing procedures. For example, composite materials often require expensive frames and mandrels in order to form a simple shape. Often, these frames must be airtight to allow a vacuum to be placed on the frame in order to bias the composite material against the interior of the frame. Additionally, these frames can require a complex resin injection system to impregnate the composite material with a resin.
Furthermore, once the composite materials are cured the attachment of multiple cured composite structures can be problematic. Currently, most complex composite structures are manufactured by assembling multiple cured composite sections. The sections are typically attached together through traditional fasteners, such as screws and bolts. This requires various fastener holes to be made into the composite materials. Each hole in the composite material severs fibers in the composite material, thus weakening the material. Other current methods of assembling cured composite structures are through post-curing adhesive bonding. However, both adhesive bonding and attaching traditional fasteners require time consuming assembly steps which are ultimately very expensive.
The expensive and complex steps for assembling a composite structure limit the widespread use of composite materials. For example, one field where the disadvantages in composite materials restrict the widespread use is in the aeronautics field. In aeronautics it is desirable for the fuselage of an aircraft to be high strength and lightweight. A lightweight fuselage increases the fuel efficiency and reduces the operating costs of an aircraft. The high strength of composite materials reduces the potential for damage to the fuselage caused by the compression and decompression of the fuselage during multiple flights.
However, the advantages provided by composite materials in fuselage construction are often not beneficial enough to overcome the disadvantages. For example, the high manufacturing costs of the fuselage can prevent the initial purchase of the aircraft, despite their lower fuel costs over a lifetime of operation. Additionally, the multiple fasteners required to assemble the various sections of cured composite structures can reduce the strength provided by the composite material. Another disadvantage is the inspection required for individual fasteners and bonding joints in aeronautic applications. Such inspections can be time-consuming and add further cost in manufacturing the fuselage.
Therefore, there is a need in the art for an inexpensive composite material manufacturing process. There is also a need in the art for an assembly process capable of co-curing the various sections of composite materials. There is a further need in the art for an assembly process capable of manufacturing complex composite structures. A need also exists for an assembly process that requires minimal fasteners and fastener holes. There is an additional need for a process that eliminated post-curing bonding steps.
There is a further need in the art for a fuselage that limits the number of fasteners present in the fuselage. There is also a need for a fuselage that may be assembled with limited assembly procedures. There is a need for a fuselage that is lightweight. A need also exists for a fuselage that limits the extent of inspection required in post manufacturing and during maintenance. A need further exists for a low-cost composite fuselage. Such a fuselage and method for manufacturing a composite structure are disclosed and claimed herein.